Actively-cooled fiber-reinforced ceramic matrix composite rocket propulsion thrust chamber and method of producing the same

ABSTRACT

An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry. The method offers the flexibility of producing internal cooling channels of either constant or continuously variable cross-sectional area, in addition to orienting the cooling channels either axially, helically or sinusoidally (e.g., undulating) with respect to the longitudinal axis of the tubular shell structure with without significant added manufacturing complication.

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This application is a continuation-in-part of application Ser.No. 09/772,108 filed Jan. 25, 2001.

FIELD OF THE INVENTION

[0002] The present invention relates to actively-cooled tubular shellstructures for high-temperature applications having a plurality ofcooling channels formed between the inner wall and outer wall of thestructure wherein the structure is fabricated from fiber-reinforcedceramic matrix composite materials and methods of producing same.

BACKGROUND OF THE INVENTION

[0003] Performance of advanced chemical rocket technology is limited,for the most part, by the availability of high-temperature structuralengineering materials. Greater performance and efficiency in liquidrocket propulsion systems can be gained by operating at highercombustion temperatures and higher working pressures resulting fromhighly energetic propellant mixture ratios (oxidizer/fuel). However,propellant mixtures are typically “tamed” to off-optimum conditions inorder not to exceed the temperature limitations of the thrust chamberand nozzle materials.

[0004] A considerable amount of heat is transferred in all designs ofrocket engines. The principle objective of high-temperature rocketdesign is to safely limit the heat transfer to the materials in criticalhot sections such as the injector, combustion chamber, throat, andnozzle. A failure would impair the satisfactory operation of the rocketpropulsion system or the flight vehicle being propelled. The walls haveto be maintained sufficiently cool so that wall temperatures do notexceed their safe allowable operating limit. Erosion, usually the resultof combined oxidation and chemical interaction with the hot combustiongases, should not damage the walls, and the walls should be capable ofwithstanding the extreme thermal shock caused by the sudden onset of ahigh heat flux from combustion ignition. The materials comprising thethrust chamber devices must also be capable of resisting the thermalstresses induced by the heat transfer and thermal gradients.

[0005] There are two general cooling methods commonly used today in thedesign of liquid propellant rocket engine devices. Those devices thatreach thermal equilibrium during operation typically operate for longdurations and are usually either actively-cooled (e.g., bipropellants)or radiation cooled (e.g., monopropellants).

[0006] Actively-cooled liquid propellant thrust chambers have provisionsfor cooling some or all of the components in contact with the hotcombustion gases, such as the chamber walls, nozzle walls and injectorfaces. A cooling jacket or cooling coil often consists of separate innerand outer walls or a bundled assembly of continuous, contoured tubes.The inner wall confines the combustion gases, and the space between theinner and outer walls serves as the coolant passage. The axial orhelical passages in liquid propellant rocket thrust chambers are oftenof complex cross-section. The nozzle throat region is usually thelocation that sustains the greatest heat transfer intensity and istherefore the most difficult to cool. For this reason the cooling jacketis often designed so that the coolant velocity is highest at thecritical regions by restricting the coolant passage cross-section and sothat the coolant enters the jacket at or near the nozzle.

[0007] Regenerative cooling is a form of active cooling and is used forengines where one of the propellant constituents is circulated throughcooling passages around the thrust chamber prior to injection andburning of the propellant in the combustion chamber. Regenerativecooling in bipropellant engines uses either the fuel or oxidizer as thecooling fluid. Therefore, the thermal energy absorbed by the coolant isnot wasted as it augments the initial energy content of the propellantprior to injection, thereby increasing the exhaust velocity andpropulsive performance.

[0008] Radiation cooling is typically used in monopropellant thrustchambers, some gas generators and for nozzle exhaust sections. Radiationcooling is a simple, lightweight cooling method, which is commonlyemployed in low-temperature rocket engines, such as hydrazine(monopropellant) spacecraft maneuver and attitude control systems, wherethe maximum chamber temperature is only about 650° C. Refractory metalssuch as molybdenum, tantalum, tungsten, and niobium have been used inradiation cooled thrust chambers requiring increased operatingtemperatures up to 1650° C. Refractory metals are, however, difficult tofabricate and some suffer from hydrogen embrittlement degradation;others oxidize readily and thus require protective surface coatings tofunction reliably; all are weight inefficient due to their very highspecific gravities. Radiation cooled thrust chambers generally have toprotrude beyond the outer skin of the flight vehicle to permitsatisfactory radiative heat rejection.

[0009] In general, ceramics have superior high-temperature strength andstiffness, and lower density than metallic materials. The principaldisadvantages of ceramics as structural materials are their low failurestrain, low fracture toughness and catastrophic brittle failurecharacteristics. Because of these inherent limitations, monolithicceramics lack the properties of reliability and durability that arenecessary for structural design acceptance. However, the emergingtechnology of fiber-reinforced ceramics, or ceramic matrix composites isone promising solution for overcoming the reliability and durabilityproblems associated with monolithic ceramics. By incorporating highstrength, relatively high modulus fibers into brittle ceramic matrices,combined high strength and high toughness composite materials can beobtained. Successfully manufactured ceramic matrix composites exhibit ahigh degree of non-linear stress-strain behavior with ultimatestrengths, failure strains and fracture toughnesses that aresubstantially greater than that of the unreinforced matrix.

[0010] In order to exploit the benefits of fiber reinforcement inbrittle ceramic matrices, it is well recognized that a relatively weakfiber/matrix interfacial bond strength is essential for preventingcatastrophic failure from propagating matrix cracks. The interface mustprovide sufficient fiber/matrix bonding for effective load transfer, butmust be weak enough to debond and slip in the wake of matrix cracking,leaving the fibers to bridge the cracks and support the far-fieldapplied load. Fiber-reinforced ceramic matrix composites with very highfiber/matrix interfacial bond strengths (usually the result of chemicalinteraction during manufacture) exhibit brittle failure characteristicssimilar to that of unreinforced monolithic ceramics by allowing matrixcracks to freely propagate directly through the reinforcing fibers.Conversely, by reducing the interfacial bond strength, the fiber andmatrix are able to debond and slip promoting the arrest and/or diversionof propagating matrix cracks at/or around the reinforcing fiber. Sincecrack inhibition/fracture toughness enhancement is the primary advantageof fiber-reinforced ceramic matrix composites, properly engineered fibercoating systems are thus key to the structural performance of thesematerials. Control of interfacial bonding characteristics between thefiber and matrix following manufacture and during service is typicallyprovided by the use of applied fiber coatings.

[0011] Fiber-reinforced ceramic matrix composites produced by thechemical vapor infiltration (CVI) process are a particularly promisingclass of engineered high-temperature structural materials, which are nowcommercially available. The principal advantage of the CVI processapproach for fabricating ceramic matrix composites as compared to othermanufacturing methods (e.g., reaction bonding, hot-pressing, meltinfiltration, or polymer impregnation/pyrolysis) is the ability toinfiltrate and densify geometrically complex, multidirectional fibrouspreforms to near-net-shape with a ceramic matrix of high purity andcontrollable stoichiometry without chemically, thermally or mechanicallydamaging the relatively fragile reinforcing fibers. In addition, becauseit is a relatively low temperature manufacturing process, high purityrefractory matrix materials can be formed (deposited) at a smallfraction of their melting temperature (˜T_(m)/4). Despite the manypossible high-temperature ceramic matrix composite material systems,however, the number of practical systems is limited by the currentlyavailable reinforcing fibers. To date, the majority of high performanceceramic matrix composites produced have focused primarily on carbon andpolymer-derived SiC (Nicalon and Hi-Nicalon) fiber reinforcement andCVI-derived SiC matrices.

[0012] Carbon fiber-reinforced silicon carbide (C/SiC) and siliconcarbide fiber-reinforced silicon carbide (SiC/SiC) composites producedby CVI have been identified by the aerospace and propulsion communities(both in the United States and in Europe) as an important enablingmaterials technology for thermostructural applications demanding highstrength and toughness at temperatures to 1650° C. The high purityCVI-SiC matrix is not readily attacked by either hydrogen-rich oroxidizing environments up to 1650° C., and resists oxidation by theformation of an adherent and protective oxide surface scale. Along withbeing chemically compatible with a variety of commercially availablerefractory fiber reinforcements, SiC possess excellent thermal shockresistance due to its combination of very high thermal conductivity(50-100 W/m/K) and low thermal expansivity (4.5 ppm/° C.). Thesethermophysical attributes are especially attractive for advanced rocketpropulsion thrust chamber and exhaust nozzle applications.

[0013] Uncooled, single-walled rocket thrust chambers and nozzlecomponents have been produced from fiber-reinforced ceramic matrixcomposites by a number of manufacturers. Material systems for theseapplications have included carbon fiber-reinforced carbon, orcarbon/carbon (C/C), C/SiC and SiC/SiC composites produced by variousmanufacturing techniques, including CVI. In all cases, these rocketpropulsion devices were fabricated as a simple, single-wall shellconstruction and passively cooled by radiation.

[0014] Until now, actively-cooled reinforced ceramic matrix compositethrust chambers have neither been designed nor produced for varioustechnical reasons, including the lack of conception of a practicaldesign approach, the perceived high level of complexity for such afiber-reinforced composite design, manufacturing difficulty, and highcost necessary to produce such a device.

SUMMARY OF THE INVENTION

[0015] In the present invention, a process is described for the designand manufacture of actively-cooled fiber-reinforced ceramic matrixcomposite tubular shell structures for high-temperature applicationswith emphasis on converging-diverging thrust chambers for liquid rocketpropulsion systems.

[0016] Tubular shell structures are defined as any open-endedthree-dimensional body with a central longitudinal axis wherein the bodyof the structure is enclosed by either curved surfaces, flat surfaces orcombinations thereof.

[0017] Fiber-reinforcement is defined as any refractory fibers, eithercontinuous or discontinuous, used for producing a fibrous preformtexture, which are capable of withstanding a use temperature of at least800° C. in an atmosphere which is thermochemically compatible with thatfiber without suffering fundamental chemical, physical or mechanicaldegradation. Examples include carbon fibers, silicon carbide fibers,silicon nitride fibers, aluminum oxide fibers, etc.

[0018] A fiber preform is a fibrous texture defined as any assemblage ofone or more reinforcing fiber types produced by weaving, braiding,filament winding, fiber placement, felting, needling, or other textilefabrication process.

[0019] Fiber preforming is a textile fabrication process by which thecollimated multifilamentary fiber bundles (tows) are placed andmaintained in a fixed position for purposes of controlling both theirorientation and content within a given volumetric space. As such, thespatial arrangement of fibers is referred to as a preform architecture.

[0020] The braided architecture is one of the simplest and lowest costfiber preforms for producing continuous fiber-reinforced axisymmetric,tubular structures. Although constructed of interlaced tows in a planar2-dimensional arrangement similar to woven fabric, braiding offers theadded flexibility of interlacing fiber in three (3) directions bothaxial and helical. The most common braid architecture consists ofhelical fibers interlaced at a prescribed bias angle and are termedbiaxial braids. Fixed axial fibers can be inserted around the mandrelcircumference in nearly any desired fraction with respect to the helical“braiders” to produce an architecture reinforced in three (3) discretedirections and are termed triaxial braids. Triaxially braided preformsoffer certain benefits over biaxial architectures, such as the abilityto tailor material isotropy or increase axial properties. In most cases,increased axial properties are gained at the cost of sacrificedcircumferential “hoop” properties. However, the optimum braidarchitecture for a given application is usually designed and selected onthe basis of the combined axial and hoop properties required. Triaxialpreforms typically yield preforms with slightly lower fiber volumefractions than that of the biaxial braided architecture. This is due tothe crimping and bunching of the fiber tows at the braid triple point,which results in increased braid layer thicknesses and larger internalvoids.

[0021] Fiber coating is defined as any refractory composition of eithercarbon, metal carbide, metal nitride, metal boride, metal silicide,metal oxide, or combinations thereof which is (are) deposited (forexample by chemical vapor infiltration) onto the refractory fiberseither before or after fiber preforming for purposes of controlling thefiber/matrix interfacial bonding characteristics in the resultantcomposite. The resultant fiber coating thus encapsulates the reinforcingfibers. Examples include pyrolytic carbon, silicon carbide, siliconnitride, boron carbide, boron nitride, etc.; either as a single-layerphase, multilayered phase or as a phase of mixed composition.

[0022] Ceramic matrix is defined as any refractory composition of eithercarbon, metal carbide, metal nitride, metal boride, metal silicide,metal oxide, or combinations thereof which is subsequently deposited(for example by chemical vapor infiltration) onto the previously coatedrefractory fibers within the fibrous preform thereby encapsulating thefibers and consolidating the preform into the resultant densifiedcomposite. The reinforcing fibers of the fibrous preform thus becomeembedded within and supported by the surrounding matrix. Examplesinclude pyrolytic carbon, silicon carbide, silicon nitride, boroncarbide, boron silicide, etc., either as a single phase, multilayeredphase or as a phase of mixed composition.

BRIEF DESCRIPTION OF THE FIGURES

[0023] FIGS. 1 (A-E) show one embodiment of an actively-cooledfiber-reinforced ceramic matrix composite tubular structure comprisingan axisymmetric rocket propulsion thrust chamber of converging-diverginggeometry.

[0024] FIGS. 2 (A-B) show a two-piece inner tooling mandrel suitable fordefining the inner wall geometry of a converging-diverging thrustchamber wherein the mandrel is split and joined at the minimumcross-section (e.g., throat) for easy removal and reusability.

[0025]FIG. 3 shows one configuration of a single axial cooling channeltooling mandrel segment suitable for defining the internal coolingchannel passage of an actively-cooled tubular structure ofconverging-diverging geometry.

[0026] FIGS. 4 (A-B) show a base cooling channel tooling mandrel ofcompound converging-diverging geometry prior to segmenting (FIG. 4A) andfollowing segmenting into a plurality of individual cooling channeltooling mandrel segments (FIG. 4B).

[0027] FIGS. 5 (A-D) show the assembly configuration of the plurality ofindividual cooling channel tooling mandrel segments disposed about theinner tooling mandrel.

[0028]FIG. 6A shows a fibrous preform overlaid onto an inner toolingmandrel of converging-diverging geometry.

[0029]FIG. 6B shows a plurality of fibrous preforms overlaid onto theindividual cooling channel tooling mandrel segments ofconverging-diverging geometry.

[0030]FIG. 6C shows a fibrous preform sub-assembly ofconverging-diverging geometry comprising a plurality of preformedcooling channel tooling mandrel segments disposed about the preformedinner tooling mandrel.

[0031]FIG. 7 shows a pair of fully completed fibrous preform assembliesconstructed in accordance with the present invention shown with theinner tooling mandrel removed for illustration.

[0032]FIG. 8 shows a pair of densified fiber-reinforced ceramic matrixcomposite rocket propulsion thrust chamber structures ofconverging-diverging geometry shown prior to the removal of theplurality of embedded cooling channel tooling mandrel segments.

[0033]FIG. 9 shows a pair of densified fiber-reinforced ceramic matrixcomposite rocket propulsion thrust chamber structures ofconverging-diverging geometry wherein the plurality of cooling channeltooling mandrel segments have been removed.

DETAILED DESCRIPTION OF THE INVENTION

[0034]FIG. 1 shows one embodiment of an actively-cooled fiber-reinforcedceramic matrix composite tubular shell structure comprising anaxisymmetric converging-diverging rocket propulsion thrust chamber 10.It can be appreciated that other tubular shell geometries (e.g.,cylindrical tubes, conical tubes, rectangular tubes, etc.) can befabricated with internal cooling channels without departing from thescope of the present invention. The basic structural arrangement of theactively-cooled thrust chamber 10 comprises three (3) fundamentalsub-structural elements: an inner combustion chamber vessel wall 11, anannular array of integral cooling channels 12 each separated byperiodically spaced radial coupling webs 13, and an outer primarypressure vessel wall 14. The inner vessel wall 11 functions as a “hot”liner in contact with and directing the flow of combustion gasesproduced during operation. The inner vessel wall 11 is preferablydesigned to be as thin as structurally possible and fabricated from amaterial possessing combined high thermal conductivity andhigh-temperature mechanical properties (e.g., strength, toughness andstiffness). A thin-wall inner shell 11 constructed of a high thermalconductivity material is preferred in order to maximize the transfer ofthermal energy from the hot inner wall 11 to the cooling fluid passedthrough the cooling channel passages 12 during operation. A materialwith high-temperature strength and toughness is preferred because of theextreme high-temperatures and pressures imposed on the inner liner 11 bythe combustion products.

[0035] In order to provide for active cooling of the thrust chamberstructure 10 shown in FIG. 1, an annular array of embedded coolingchannels 12 is provided. These channels 12 serve to contain andtransport pressurized cooling fluid through the interior of the thrustchamber structure 10. Because it is desirable for the thermallyeffective surfaces of the cooling channels 12 to be in intimate contactwith the hot inner wall 11 for maximum cooling efficiency, atrapezoidal-shaped cooling channel passage 12 cross-sectional geometryas shown in FIG. 1 is preferred in order to maximize the contact surfaceadjoining the inner wall 11 and the cooling fluid. It can be appreciatedthat other cooling channel geometries (e.g., circular) can be utilizedfor fabrication of the cooling channels without departing from the scopeof the present invention. However, the trapezoidal-shaped coolingchannel geometry 12 is preferable for the present application over thatof circular based on: (1) providing the maximum wetted surface area perunit channel packing, resulting in the most efficient convective heattransfer (e.g., cooling) characteristics; (2) the ability tocontinuously vary the cooling channel cross-sectional area so as topassively control local coolant flow rates for optimal thermal control;(3) providing the greatest structural integrity via direct, linear loadpaths from the inner vessel wall 11 to the outer vessel wall 14; (4)minimizing channel-to-shell nesting voids 15, thereby minimizinginterlaminar stresses and diminished heat transfer efficiency; (5) theability to easily integrate the inner shell wall 11, outer shell wall 14and cooling channel 12 components into a single monocoque assembly 10;and (6) the relative ease of manufacturing the geometrically-complexcoolant channel tooling mandrels without the need for an elaborate andexpensive multi-axis, numerically-controlled milling machine operation.

[0036] A plurality of discrete cooling channels 12 is preferred over anopen annulus, dual-shell (e.g., cooling jacket) cooled structure designconfiguration. The array of individual cooling channel passages 12 aidsto straighten the flow of the cooling fluid passed therethrough, therebyminimizing undesirable fluid mechanical effects (e.g., rotational flow),which can compromise the convective heat transfer efficiency of the heatexchanger.

[0037] The cooling channels 12 are separated from each other byperiodically spaced radial web structures 13. The plurality of radialwebs 13 serve to structurally couple the inner 11 and outer 14 vesselwalls into a single integrated monocoque assembly 10. The webs 13provide a direct radial load path from inner wall 11 to outer wall 14,which serves to increase the weight-specific structural efficiency ofthe overall thrust chamber vessel 10. The webs 13 also provide discreteradial thermal conduction pathways (e.g., radial cooling fins), whichserve to increase the effectiveness of the convective heat exchanger.

[0038] The actively-cooled structure 10 is fabricated from afiber-reinforced ceramic matrix composite material specially designedfor high-temperature structural applications, such as rocket propulsionthrust chambers, burner tubes, heat exchangers, and other severeenvironment thermostructural components. The ceramic matrix compositematerial comprises a fibrous preform composed of an assemblage ofrefractory fibers, a fiber coating material encapsulating the refractoryfibers, and a ceramic matrix material that consolidates the coatedfibrous preform into a densified composite.

[0039] A first step in producing a near net-shape actively-cooledfiber-reinforced ceramic matrix composite tubular shell structure 10 inaccordance with the present invention is the fabrication of toolingmandrels. The tooling mandrels serve to provide structural support andgeometric definition of the fibrous preform, which will ultimatelydefine the geometry of the resultant composite structure following fibercoating, matrix densification and mandrel removal.

[0040] Suitable tooling mandrel materials should be selected on thebasis of several considerations: (1) thermochemical stability in thefiber coating and matrix densification processes; (2) thermochemical andthermomechanical compatibility with the preform reinforcing fiber; (3)thermophysical stability for maintaining desired component geometryduring and following materials processing; (4) adequate structuralintegrity for handling the fibrous preform textile fabrication processesand subsequent preform assembly; (5) ease of removal; and (6) cost.Examples of suitable materials include graphite, molybdenum andstainless steel.

[0041] As a generalized example, the inner mold line (e.g., innersurface) of an axisymmetric converging-diverging ceramic matrixcomposite thrust chamber 10 as shown in FIG. 1 is produced and definedby an inner tooling mandrel. Because of its axisymmetry, this innertooling mandrel can be fabricated using standard lathe turningtechniques. Lathe machining of the converging-diverging chamber geometryfrom a suitable tooling material can be accomplished using a simpleprofile tracer template, numerical control, or other machiningtechniques well known in the art. FIG. 2 shows an example of aconverging-diverging split inner tooling mandrel 20 suitable for use infabricating the inner shell wall 11 of an actively-cooled ceramic matrixcomposite rocket propulsion thrust chamber 10.

[0042] For the example of a converging-diverging geometry, the innertooling mandrel 20 can also be fabricated from two (2) segments, whichcan be split at the minimum cross-section 21 (e.g., throat) andmechanically joined by an axial bolted fastener to facilitate easyremoval and reusability of the inner tooling mandrel 20. A simpleactively-cooled cylindrical tube geometry on the other hand can beeasily defined by a simple, single-piece cylindrical inner toolingmandrel of constant diameter along its length. Regardless of the finaltubular geometry desired, excess length at either end of the innertooling mandrel 20 may be desirable to accommodate certain fiberpreforming methods, such as braiding, and subsequent fibrous preformassembly as will be described. In this event, any irregular and/orexcess material at the ends of the densified composite structure can betrimmed and discarded following completed fabrication.

[0043] Following construction of an inner tooling mandrel 20 fordefining the inner wall 11 geometry of the ceramic matrix compositetubular shell structure 10 to be produced, a fibrous preform ofpredetermined architecture and thickness is constructed onto the innertooling mandrel 20 by a suitable textile fabrication process, such asbraiding. The inner fiber preform forms the inner shell wall 11sub-structure of the actively-cooled composite structure 10; thus, thepreform architecture and preform thickness should be chosen with thisdesign consideration in mind.

[0044] For increased resistance to damage or distortion sustained bysubsequent handling, the fiber preform can be restrained onto the innermandrel by hoop wrapping the excess ends with a suitable reinforced tapeor string, as these ends may eventually be trimmed away and discarded.

[0045] The inner mold line of the array of annularly arranged coolingchannel passages 12 within the composite structure 10 are produced anddefined by individual tooling mandrel segments. The number of coolingchannels passages 12, their corresponding orientation (e.g., axial,helical or sinusoidal), and cross-sectional area properties are selectedbased on satisfying the combined structural, thermal and fluidmechanical requirements of the actively-cooled structure as determinedfrom design.

[0046] Once a suitable design configuration for the cooling channels hasbeen selected, individual cooling channel mandrel segments are thenfabricated in accordance to this selection. FIG. 3A shows an example ofa single axial cooling channel mandrel segment 30 suitable for use infabricating an actively-cooled fiber-reinforced ceramic matrix compositethrust chamber 10 of general converging-diverging geometry. Fabricationof a plurality of individual tooling mandrels 30 having the complexconvoluted geometry shown in FIG. 3 would normally require sophisticated(e.g., numerically controlled) multi-axis mill machining. Fabrication ofhelical or sinusoidally undulating channel tooling mandrels would beeven more complex. The present invention, however, describes a greatlysimplified manufacturing technique for fabricating the geometricallycomplex cooling channel mandrel segments 30 necessary for producing anactively-cooled ceramic matrix composite converging-diverging thrustchamber 10 or other enclosed tubular structure.

[0047] In the present invention, a “base” cooling channel toolingmandrel corresponding to the selected geometry of the cooling channelpassages 12 to be produced within the structure is fabricated from asuitable tooling material using standard lathe turning and mill or sawcutting techniques. Selection of a suitable tooling material for thefabrication of the base cooling channel mandrel is important, aspreviously described, and must also take into consideration the processto be used for subsequent removal of the individual tooling mandrelsegments, which may become embedded or intimately attached following thematrix densification of the actively-cooled ceramic matrix compositestructure. As discussed below, graphite is a suitable choice forfabrication of the base cooling channel mandrel as it may besubsequently removed using conventional techniques (e.g., bead blasting)without damaging the surrounding ceramic matrix composite structure.Other tooling materials may also be appropriate and are discussed below.

[0048] The base cooling channel tooling mandrel for an axisymmetricconverging-diverging cylindrical geometry (e.g., actively-cooled thrustchamber) is fabricated by machining the outer diameter of the selectedtooling material with a desired profile, which does not necessarily haveto correspond to the profile of the inner mandrel 20 previouslydescribed. Following machining of the outer profile, an inner profile isthen produced onto the mandrel which directly corresponds to thegeometric profile of the inner mandrel 20. The difference between theinner and outer diameters of the base cooling channel mandrel ultimatelydefines the cooling channel passage cross-sectional area as will bedescribed below. FIG. 4A shows a suitable base cooling channel mandrel40 for producing the individual cooling channel mandrels 30 of aconverging-diverging thrust chamber structure 10 as described above.With reference to FIG. 4A, the resultant base cooling channel mandrel 40for a converging-diverging thrust chamber 10 comprises a hollow tubularshell structure having compound inner and outer converging-divergingprofiles. As previously described, the inner and outer profiles of thebase cooling channel mandrel 40 can be independent of one another andwill thus form the desired cross-sectional geometry of the internalcooling passages 12. The inner profile of the base mandrel 40 isdesigned to geometrically conform to the profile of the inner mandrel 20with prescribed radial accommodation (e.g., relief or gap) for thethickness of fiber preforms, which are ultimately produced over therespective inner mandrel 20 and individual cooling channel mandrels 30following their fabrication. Thus, the radial dimensions of the innermandrel 20 and the internal contour of the base mandrel 40 differ onlyby the fiber preform thickness desired along the length of therespective tools, and ultimately in the completed composite structure.

[0049] Following fabrication of the inner and outer profiles, thecompound converging-diverging base mandrel 40 as shown in FIG. 4A isthen cut into the desired axial, helical or sinusoidal segments usingconventional cutting saw or vertical milling machining techniques. Theresultant cut segments thus produce the individual cooling channelmandrels 30 which are used to form the internal cooling channel passages12 of the actively-cooled ceramic matrix composite structure 10. FIG. 4Bshows a converging-diverging base cooling channel mandrel 40 that hasbeen cut axially into individual cooling channel tooling mandrelsegments 30.

[0050] The cutting process for fabricating the individual coolingchannel mandrel segments 30 from the base mandrel 40 is preferablyperformed by using a prescribed kerf, or cut width in order to controlthe desired thickness of the fibrous preform that, as described below,will eventually be produced onto and formed by the individual coolingchannel tooling mandrel segments 30. The appropriate kerf width forcutting the cooling channel mandrel segments 30 from the base mandrel 40should be approximately twice the desired fibrous preform thickness tobe produced on each channel mandrel. This is to accommodate thecumulative preform thickness build-up from adjoining cooling channelmandrel segments upon their assembly and compaction. The cumulativepreform thickness between each adjoining cooling channel fiber preformfollowing their annular assembly forms the array of radial coupling webs13 of the actively-cooled ceramic matrix composite structure 10.

[0051] For axially oriented channels 12, the individual cooling channelmandrel segments 30 are produced simply by a systematic process oflongitudinal cutting followed by rotational indexing of the base mandrel40 at discrete angular intervals. The angular indexing interval(360-degrees/n) is selected based on the number of like cooling channels(n) desired. This systematic process of cutting and indexing is repeateduntil the entire base mandrel shell 40 is cut into the desired number ofequal segments 30. FIG. 4B shows an example of a plurality ofaxially-oriented cooling channel mandrel segments 30 produced from asingle compound converging-diverging base cooling channel mandrel 40 asis shown in FIG. 4A.

[0052] For helically oriented channels, the individual cooling channelmandrel segments may be produced by simultaneously rotating the basecooling channel mandrel 40 about its longitudinal axis while cutting.The rate of angular rotation combined with the rate of longitudinalcutting is synchronized to produce the desired helical cut path.Following each helical cut, the base cooling channel mandrel 40 isrotationally indexed as previously described and the cutting process isrepeated in manner similar to that of axially aligned cooling channelmandrel segments 30.

[0053] For sinusoidally undulating channels, the individual coolingchannel mandrel segments are produced by simultaneously undulating thebase cooling channel mandrel 40 about its longitudinal axis in a cyclicrotational or back-and-forth motion while cutting. The rate of therotational reversal while cutting ultimately defines the amplitude ofthe undulating individual cooling channel mandrel segments produced. Therate of longitudinal cutting is thus synchronized to produce the desiredsinusoidal period, or wavelength. Following each sinusoidal cut, thebase cooling channel mandrel 40 is rotationally indexed and the cuttingprocess is repeated in a manner similar to that of axially alignedcooling channel mandrel segments 30.

[0054] To aid in ease of fabricating the cooling channel tooling mandrelsegments, it is best not to fully cut the entire length of the basemandrel 40 during segmenting of the individual cooling channel mandrels30 as the remaining uncut base mandrel 40 becomes flexible and awkwardto handle and accurately index during machining. By retaining shortuncut lengths at both ends of the base mandrel 40 during cutting, thebase mandrel 40 remains circumferentially rigid and adequatelystabilized by the remaining uncut portion of the attached individualcooling channel mandrel segments 30. After completion of cutting allcooling channel mandrel segments 30, the uncut ends of the base coolingchannel mandrel 40 are then cut perpendicular to the longitudinal axisof the base mandrel 40 and discarded, thereby freeing all the individualcooling channel mandrel segments 30. The final length of the individualcooling channel mandrel segments 30 is therefore shortened by the lengthof the discarded trim material. Thus, the base mandrel 40 can be madelonger to accommodate this subsequent end trimming method. As-machinedsharp edges on the individual cooling channel mandrel segments 30 can beremoved by chamfering if necessary. Chamfering aids in minimizingpotential fiber damage from abrasion during fiber preform fabrication.Although beneficial for fiber preforming, excessive chamfering can causeexcessive preform nesting assembly voids 15 in the resultant compositestructure 10.

[0055]FIG. 5 shows the assembly form of individual cooling channelmandrel segments 30 disposed around an inner tooling mandrel 20 forproducing an actively-cooled ceramic matrix composite thrust chamber 10of converging-diverging geometry in accordance with the presentinvention (axial cooling channel design shown without fibrous preformfor illustration). As shown in FIG. 5, a predetermined machining kerfhas been chosen to allow for circumferential relief (e.g., gaps) betweeneach adjacent cooling channel mandrel segment 30 in order to form theradial coupling webs 13. Radial relief between the inner surface of theindividual cooling channel mandrel segments 30 and the adjacent innertooling mandrel 20 is also provided in order to form the inner vesselwall 11. It is important that these relief features be designed inaccordance with the selected thickness of the fibrous preformssubsequently produced onto the individual cooling channel mandrelsegments 30 and the inner tooling mandrel 20 for proper assembly. Forexample, a kerf of greater than twice the preform thickness on theindividual cooling channel mandrel segments 30 will result in producinga composite with excessive nesting voids 15, whereas a kerf of less thantwice the preform thickness on the individual cooling channel mandrelsegments 30 will result in undesirable assembly interference.Significant assembly interference can result in producing a compositewith either insufficient inner wall fiber volume fraction ordelamination between the inner wall 11 and adjoining cooling channels12, which can significantly diminish both the structural and thermalperformance of the resultant actively-cooled ceramic matrix compositestructure 10.

[0056] Following cooling channel mandrel fabrication, a fibrous preformof predetermined architecture and thickness is then produced onto eachof the individual mandrel segments 30 by a suitable preform fabricationprocess, such as braiding.

[0057] Having an overlaid fiber preform, the individual cooling channeltooling mandrel segments 30 are then assembled around the perimeter ofthe inner tool mandrel 20 in an annular array. As described previously,the inner tooling mandrel 20 has already been overlaid with its owninner fibrous preform. The individual cooling channel mandrel segments30 are preferably held into position and compacted against the innerfibrous preform radially and against themselves circumferentially toobtain the desired preform thickness in the inner wall 11 and radialwebs 13, respectively. Control of these thicknesses via tooling design(e.g., preform relief), fiber preforming and preform assembly ultimatelycontrols the fiber volume and nesting void fraction within the resultantcomposite structure. Restraint of the compacted channel fiber preformsegments against the inner fiber preform may be accomplished by tightlyhoop wrapping the ends of the channel mandrel segments with a suitablereinforced tape or string. It is preferable to restrain thetooling/preform sub-assembly in the end region of excess length as theseends can be trimmed away and discarded following preform consolidationand matrix densification processing.

[0058]FIG. 6A shows the inner tooling mandrel 20 having an overlaidfibrous preform 60, which forms the reinforcement of the inner wall 11of the ceramic matrix composite structure 10. FIG. 6B shows a number ofcooling channel mandrel segments 30 having overlaid fibrous preforms 61which form the reinforcement of the individual cooling channels 12 ofthe ceramic matrix composite structure 10. FIG. 6C shows the assembly ofindividual cooling channel mandrel segments 30 having overlaid fibrouspreforms 61 disposed around the inner tooling mandrel 20 and innerfibrous preform 60. Compacting and securing the individual coolingchannel fibrous preform segments 61 against the inner fibrous preform 60as a fibrous preform sub-assembly 62 is provided by hoop-wrapping 63 theexcess ends of the individual cooling channel mandrel segments 30.

[0059] Once the fiber preform sub-assembly 62 has been fabricated, it isthen overlaid with an outer fibrous preform by a suitable textilefabrication process, such as braiding. The outer fibrous preform willthus conform to the geometry of the underlying tooling/preformsub-assembly 62. FIG. 7 shows a completed converging-diverging thrustchamber tooling/fibrous preform assembly 70 (inner tooling mandrel 20removed for illustration) ready for subsequent fiber coating and matrixdensification. For ease of handling in preparation for subsequent fibercoating and/or matrix consolidation processing, the entire fibrouspreform assembly 70 can also be restrained by hoop wrapping the excessends with a suitable reinforced tape or string 72, again, as these endscan eventually be trimmed away and discarded. The outer fibrous preform71 forms the outer wall 14 of the actively-cooled ceramic matrixcomposite structure 10 produced. Thus, the fabrication of the fibrouspreform 70 corresponding to the desired actively-cooled tubularstructure 10 to be produced is completed following the deposit of theouter fiber preform 71 onto the underlying fibrous preform sub-assembly62.

[0060] Following fabrication of the fibrous preform assemblage 70, afiber coating or fiber coating system is deposited (for example bychemical vapor infiltration) onto the entire tooling/preform assembly70. If the reinforcing fibers have been coated prior to preforming, thisprocessing step may not be necessary. As previously described, the fibercoating serves to control the desired fiber/matrix interfacialcompliance and bonding characteristics in the resultant ceramic matrixcomposite structure produced. Examples of suitable fiber-coatingmaterials include pyrolytic carbon, silicon carbide, silicon nitride,boron carbide, boron nitride, etc., either as a single layer phase,multilayered phase or as a phase of mixed composition.

[0061] Following the application of the fiber coating or fiber coatingsystem, a ceramic matrix is deposited (for example by chemical vaporinfiltration) onto the entire tooling/preform assembly 70. The processof matrix deposition using chemical vapor infiltration is well known inthe art and is also described in U.S. Pat. No. 5,455,106 and Applicant'sco-pending patent application having Attorney Docket Number 260/175, thedisclosures of which are incorporated herein by reference in theirentirety. It can be appreciated that other matrix processing techniques(e.g., polymer impregnation/pyrolysis, melt infiltration, etc.) can beutilized for densification of the fibrous preform assembly 70 withoutdeparting from the scope of the present invention. Depending on themethod of matrix deposition processing employed, consolidation of thefibrous preform assembly 70 can be performed in two sequential steps.The first processing step involves the application of a small volumefraction of matrix to stiffen or rigidize the fibrous preform assembly70. A properly rigidized fiber preform assembly 70 lacks flexibility andis sufficiently stiff enough to be handled and further processed withoutthe aid of external fixturing and/or tooling for stabilizing andmaintaining the desired component geometry. The amount of depositedmatrix necessary to adequately rigidize the preform assembly 70 isdependent on the matrix composition and processing method employed.Following initial rigidization, the inner tooling mandrel 20 may then beremoved to enable consolidation to commence on the freestanding fibrouspreform assembly 70 as shown in FIG. 7. Early removal of the innertooling mandrel 20 following initial rigidization facilitates easymandrel extraction in addition to mandrel reusability and enhancedmatrix densification processing economics. However, for complexgeometries (e.g., converging-diverging) and/or helical or sinusoidalcooling channel designs, the individual cooling channel mandrel segments30 must remain captive within the fibrous preform assembly 70 due totheir convoluted geometry. Should straight axial cooling channel mandrelsegments be chosen, such as that suitable for actively-cooled rightcylindrical or tapered cylindrical tubes, these mandrels may beeffectively removed at this step, as they may not be geometricallycaptured within the preform. Thus, removal of the individual coolingchannel mandrel segments is dependent on the geometry of the coolingchannel mandrels chosen which, is dependant on the application for whichthe actively-cooled ceramic matrix composite structure is intended.

[0062] The second processing step following the removal of the innermandrel 20 is that of preform consolidation. Consolidation of the rigid,yet still highly porous preform assembly 70 with a ceramic matrix isthus performed until a zero or near-zero residual open porosity level isattained in the resultant composite. The reinforcing fibers of the innerfibrous preform 60, outer fibrous preform 71, and cooling channelfibrous preforms 61 thus become embedded within, supported by and fusedto the surrounding matrix. Examples of suitable matrix materials includepyrolytic carbon, silicon carbide, silicon nitride, boron carbide, boronnitride, etc., either as a single-layer phase, multilayered phase or asa phase of mixed composition.

[0063] Following matrix densification processing, the densifiedactively-cooled ceramic matrix composite structure is prepared foreither rough or final detailed machining and cooling channel toolingmandrel removal. Rough machining involves removing the excess materiallength used initially for assembling and restraining the fibrous preformfrom both ends of the actively-cooled ceramic matrix composite tubularstructure. Rough machining/trimming can be performed using standarddiamond cutting saw machining techniques. The trim end remnants are thusremoved from the densified ceramic matrix composite structure anddiscarded. FIG. 8 shows a near-densified ceramic matrix composite thrustchamber structure of converging-diverging geometry which has undergonerough machining. Cutting off the ends of the actively-cooled ceramicmatrix composite structure, as shown in FIG. 8, exposes the ends of theembedded cooling channel mandrel segments 30, which is necessary fortheir removal in the case of cooling channels having convoluted geometry(e.g., converging-diverging).

[0064] The method of removing the embedded cooling channel mandrelssegments 30 is dependent on the tooling material employed forfabrication of the base cooling channel mandrel 40. Graphite coolingchannel mandrels 30 can be removed by either oxidation or grit blasting.For composite materials not composed of carbon or other low oxidationthreshold constituents (e.g., reinforcing fiber, fiber coating ormatrix) oxidative combustion of graphite mandrels is performed bythermally conditioning the composite structure in a furnace heated to atleast 425° C. in an ambient (air) environment. The rate of mandrelremoval can be accelerated with increased conditioning temperaturesand/or increased oxygen partial pressures. Conditioning temperatures andoxidizing atmospheres should be maintained well under that necessary tocause thermochemical (e.g., oxidation) or thermomechanical degradationof the ceramic matrix composite material.

[0065] For ceramic matrix composite materials of sufficient erosionresistance (e.g., high hardness and strength), grit blasting is a viablealternative to oxidative combustion for rapid removal of graphitecooling channel mandrels 30. Because of its combination of low hardnessand strength, bulk graphite exhibits very low resistance to hardparticle erosion. Blast media and delivery pressures are selected basedon providing the most aggressive erosive removal of the mandrels withoutimparting erosive damage or compositional contamination to the adjacentcomposite material. Small diameter orifice nozzles combined with nozzleextensions can be used to enable access down the otherwise long, slenderand convoluted channel passages 12.

[0066] Metallic cooling channel mandrels 30 can be effectively removedby acid digestion. For ceramic matrix composite materials composed ofconstituents which are chemically compatible with suitable etching acids(e.g., nitric acid), dissolution of metallic cooling channel mandrelsegments 30 is performed by submerging the composite structure into anacid bath. The rate of mandrel removal can be accelerated with increasedacid bath temperatures and/or increased acid concentrations oraggressive acid mixtures. For very long and slender channel passages 12,cyclic pressure or vacuum pulsing of a digestive bath system aids in theaccelerated removal of high aspect ratio cooling channel mandrelsegments 30.

[0067] Following removal of the embedded cooling channel mandrels 30,final detailed machining is performed on the densified ceramic matrixcomposite structure to satisfy specific geometric and dimensionaltolerance requirements (e.g., parallelism and flatness of the free ends,chamfering or beveling of sharp edges, and/or lathe turning of the innerand/or outer diameter(s) to incorporate provisional details for coolantmanifold integration, etc.). Detailed machining is typically performedusing standard diamond grinding techniques.

[0068] Following removal of the individual cooling channel mandrelsegments 30 and final machining, the actively-cooled ceramic matrixcomposite structure 10 is preferably processed with a final matrix sealcoating. Seal-coat processing as a final step in the fabrication of theactively-cooled ceramic matrix composite structure serves to over-coatthe freshly exposed interior walls of the cooling channel passages 12 aswell as providing an environmentally protective coating on free-edgesurfaces freshly exposed from cutting and final machining. FIG. 9 showsa completed actively-cooled fiber-reinforced ceramic matrix compositethrust chamber structure 10 of converging-diverging geometry.

EXAMPLE

[0069] The present invention will now be described with reference to aspecific example comprising the fabrication of an actively-cooledfiber-reinforced ceramic matrix composite thrust chamber ofconverging-diverging geometry using the methods previously described.For this example, a simple axially oriented cooling channel design isselected for purposes of demonstrating the manufacturing methods of thecurrent invention. The example thrust chamber is produced with a 50.8 mmdiameter cylindrical combustion chamber, a 25.4 mm throat, a 101.6 mmdiameter conical exhaust (expansion) nozzle, and is 203.2 mm in overalllength. The convergent and divergent sections are of simple conicalgeometry, both with a half angle of 20-degrees with respect to thelongitudinal axis. The ceramic matrix composite structure is designed tosafely sustain combustion chamber pressures and coolant deliverypressures of up to 7.5 Mpa and 11.2 Mpa, respectively. The axial heatexchanger circuit is designed to reduce the combustion chamber innerwall temperature from over 3,600° C. to less than 1,650° C. usingcryogenic liquid hydrogen coolant delivered at a temperature of −220° C.and flow rate of 0.45 kg/sec.

[0070] The material system selected for demonstrating theactively-cooled thrust chamber design and manufacturing methodology is abraided Hi-Nicalon silicon carbide (SiC) fiber-reinforced CVI-SiCceramic matrix composite (SiC/SiC). The SiC/SiC composite materialsystem is selected for its high thru-thickness thermal conductivity,negligible thru-the-thickness permeability, and inherent oxidationresistance. Continuous Hi-Nicalon SiC fiber is selected as thestructural reinforcement for its excellent thermal expansioncompatibility with the CVI-derived SiC matrix and adequatehigh-temperature mechanical properties. A braided perform architectureis selected because of its amenability for producing high-performanceaxisymmetric tubular structures and tailorable in-plane (e.g., axial andcircumferential) mechanical properties.

[0071] Tooling mandrels (e.g., inner mandrel and cooling channel mandrelsegments) for the example thrust chamber are fabricated from fine-grain,high-density bulk graphite. Bulk graphite is selected as the toolingmaterial for its outstanding compatibility with the CVI-based PyC andSiC processes and low cost. The example thrust chamber design requireseighteen (18) individual channel mandrel segments, each sectioned at20-degree intervals. All tooling mandrels are fabricated with adequateadditional trim length for preform fabrication, assembly, compaction,and restraint for reasons previously described.

[0072] For this example, a biaxially braided Hi-Nicalon SiC fiberpreform is selected for its combination of high volumetric loading offiber reinforcement, excellent hoop properties and overall designsimplicity. The fibrous preform is produced by braiding four (4)fabric-like layers in succession onto the inner split mandrel and four(4) fabric-like layers onto each of the eighteen (18) cooling channelmandrel segments. The selected braid bias (helix) angle of the fiberreinforcement in each preform layer is approximately ±60-degree withrespect to the thrust chamber longitudinal axis. The ±60-degree fiberarchitecture is selected for its near-optimum balance of axial and hoopmechanical properties desired for axisymmetric composite pressure vesselstructures.

[0073] Following the fiber preforming process, the eighteen (18) channelsegments are disposed around and assembled onto the inner mandrel in anannular array. The assemblage is radially compacted and secured using atight wrap of heavy-duty carbon yarn. FIG. 6 shows a photograph of thebraided inner mandrel, nine (9) of the eighteen (18) required coolingchannel mandrel segments and the resultant sub-assembly of therespective fibrous preforms. The entire tooling/preform sub-assemblageis then overlaid with six (6) ±60-degree braid layers, therebycompleting the actively-cooled thrust chamber fibrous preform assembly.

[0074] The initial CVI processing step is the application of a suitablefiber coating onto the fibrous preform assembly. As previouslydiscussed, a fiber coating is required to impart the necessaryfiber/matrix interfacial mechanical characteristics (e.g., lowinterfacial shear strength) to promote high strength and toughness inthe resulting composite. The selected fiber coating for this example isa two-layer, duplex coating system comprised of a first ˜0.4 μm debondlayer of pyrolytic carbon (PyC) followed by a second ˜0.6 μmoxidation-barrier layer of boron carbide (B₄C) produced by CVI methods.

[0075] The CVI-derived PyC fiber coating is produced in ahigh-temperature, low-pressure chemical vapor infiltration (CVI) reactorby the thermal decomposition of a hydrocarbon-containing gas in thepresence of hydrogen according to the following chemical reaction:

1/nC_(n)H_(m)+αH₂→C+(m/2n+α)H₂,

[0076] where C_(n)H_(m) is the gaseous hydrocarbon reactant (e.g.,methane, propane, propylene etc.) and α is defined as the molar ratio ofH₂ to C_(n)H_(m). At deposition temperatures between 1000-1400° C., thedeposit is typically smooth laminar PyC, which has a hexagonal structureand, depending on the deposition temperature, has a density rangingbetween 1.8-2.0 g/cm³.

[0077] The next processing step is the application of the outer B₄Coxidation-barrier coating produced by the CVI co-deposition of boron (B)from boron trichloride (BCl₃) and carbon (C) from a suitable hydrocarbongas using hydrogen as a carrier according to the following chemicalreaction:

4BCl₃+1/nC_(n)H_(m)+αH₂→B₄C+12HCl+(m/2n+α−6)H₂,

[0078] where 4n is the molar ratio of BCl₃ to C_(n)H_(m), which iscontrolled to produce a near-stoichiometric B₄C coating composition at agiven processing condition. At deposition temperatures between1000-1300° C., the deposit is typically crystalline B₄C, which has arhombohedrel structure and a theoretical density of 2.52 g/cm³.

[0079] The next CVI processing step is the rigidization of the fibrouspreform with a ceramic matrix. The selected matrix constituent for thisexample is SiC for reasons previously discussed. The CVI SiC matrix isproduced by the thermal decomposition of vaporized methyltrichlorosilane(MTS) using hydrogen as a carrier gas at elevated temperature andreduced pressure according to the following chemical reaction:

CH₃SiCl₃+αH₂→SiC+3HCl+αH₂,

[0080] where α is defined as the molar ratio of H₂ to CH₃SiCl₃. Atdeposition temperatures between 900-1300° C., the deposit is typicallycrystalline beta-SiC, which has a cubic structure and a theoreticaldensity of 3.21 g/cm³.

[0081] Following the initial deposition of the SiC matrix, the rigidizedfiber preform is prepared for removal of the inner split mandrel priorto complete matrix densification processing. The CVI-derived SiC matrixtypically bridges between the fibrous preform in physical contact withthe adjacent inner tooling mandrel. This results in a slight, but easilyfrangible bonding of the fibrous preform to the adjacent mandrel.Removal of the split mandrel is performed by first removing the threadedaxial tie bolt from the inner mandrel assembly. The SiC coating bond isthen breached from the surrounding preform and inner mandrel and theopposing tools are then extracted from their respective ends of thepreform assembly. FIG. 7 shows a photograph of the rigidized thrustchamber preform with the inner split mandrel removed.

[0082] Consolidation processing of the free-standing fibrous preformwith the SiC matrix is then performed until a near-zero residual openporosity level is attained in the composite.

[0083] After achieving near-final densification, rough machining isperformed by trimming the excess material length from both ends of theceramic matrix composite thrust chamber component using a diamond-bladedcutting saw. Removal of the excess material exposes the ends of theindividual graphite cooling channel mandrel segments, which remaincaptive within the dense composite structure. FIG. 8 shows a photographof the densified ceramic matrix composite thrust chamber followingremoval of the trim ends.

[0084] Since the selected SiC/SiC composite is much harder and far moreerosion-resistant than the selected graphite tooling material, theembedded cooling channel mandrel segments are easily removed by gritblasting methods. Commercially available borosilicate glass bead mediaprovides adequate erosive qualities for removing the embedded graphitemandrels without inducing damage or contamination to the surroundingceramic matrix composite material. Media delivery pressures of up to 0.7MPa are satisfactory for the rapid removal of the graphite mandrelswithout damaging the underlying ceramic matrix composite material. Gritblasting is best performed using a long, slender tubular nozzleextension to enable unobstructed access through the narrow convolutedcooling channel passages of the structure.

[0085] After complete removal of the graphite channel tooling mandrels,the ceramic matrix composite thrust chamber is further processed with anadditional SiC matrix infiltration/densification step. This processingstep serves, in part, to densify and coat the freshly exposed interiorwalls of the cooling channel passages. Matrix processing is thusperformed until a zero residual open porosity level is attained in theresultant composite. A zero open porosity level in the compositerepresents the maximum degree of densification achievable by CVIprocessing methods, as the chemical reactants no longer have diffusionalaccess to the interior of the material via open porosity.

[0086] Detailed machining is then performed on the fully densifiedcomposite thrust chamber as necessary to satisfy specific geometric anddimensional tolerance requirements. Such requirements may includeparallelism and flatness of the free ends, chamfering or beveling ofsharp edges, and/or turning of the inner and/or outer diameter(s) toincorporate provisional details for external coolant manifoldintegration.

[0087] The final processing step in the manufacture of theactively-cooled ceramic matrix composite thrust chamber is that of CVISiC “seal coating”. Seal coat processing as a final step in thefabrication of the composite thrust chamber serves to provide anenvironmentally-protective SiC over-coat to all surfaces exposed fromdetailed machining, as previously described. FIG. 9. shows a photographof a fully completed actively-cooled SiC/SiC composite thrust chamberproduced by the manufacturing methods described herein.

1. A ceramic matrix composite tubular shell structure having an innerwall, an outer wall and a plurality of cooling channels formed betweensaid inner wall and said outer wall wherein said ceramic matrixcomposite tubular shell structure comprises: a. a fibrous preform ofrefractory fibers; b. a fiber coating which fully encapsulates therefractory fibers of said fibrous preform; and c. a ceramic matrixmaterial which fully encapsulates and consolidates the coated fibers ofthe said fibrous preform into a densified composite material.
 2. Theceramic matrix composite tubular shell structure recited in claim 1,wherein the refractory fibers of said fibrous preform are selected froma group comprising continuous and discontinuous high-temperature fibers.3. The ceramic matrix composite tubular shell structure recited in claim1, wherein the refractory fibers of said fibrous preform are selectedfrom a group comprising carbon, silicon carbide, aluminum oxide, andother fibers capable of withstanding temperatures in excess of 800° C.4. The ceramic matrix composite tubular shell structure recited in claim1, wherein said fibrous preform is produced by a textile fabricationprocess selected from a group of fabrication processes comprisingweaving, braiding, knitting, filament winding, felting, and needling. 5.The ceramic matrix composite tubular shell structure recited in claim 1,wherein said fiber coating is a material having a thickness of 0.05-5.0micrometers and is selected from a group of materials comprising carbon,silicon carbide, boron carbide, tantalum carbide, hafnium carbide,zirconium carbide, silicon nitride, boron nitride, tantalum nitride,hafnium nitride, zirconium nitride, titanium nitride, aluminum nitride,silicon boride, tantalum boride, hafnium boride, zirconium boride,titanium boride, zirconium silicide, titanium silicide, molybdenumsilicide, aluminum oxide, silicon oxide, boron oxide, tantalum oxide,hafnium oxide, zirconium oxide, and titanium oxide.
 6. The ceramicmatrix composite tubular shell structure recited in claim 1, whereinsaid fiber coating comprises a single-layer phase of uniform materialcomposition.
 7. The ceramic matrix composite tubular shell structurerecited in claim 1, wherein said fiber coating comprises a multilayeredphase including two or more alternating coating layers having two ormore fiber coating material compositions.
 8. The ceramic matrixcomposite tubular shell structure recited in claim 1, wherein said fibercoating comprises a single-layer phase of mixed material composition. 9.The ceramic matrix composite tubular shell structure recited in claim 1,wherein said ceramic matrix material is selected from a group ofmaterials comprising carbon, silicon carbide, boron carbide, tantalumcarbide, hafnium carbide, zirconium carbide, silicon nitride, boronnitride, tantalum nitride, hafnium nitride, zirconium nitride, titaniumnitride, aluminum nitride, silicon boride, tantalum boride, hafniumboride, zirconium boride, titanium boride, zirconium silicide, titaniumsilicide, molybdenum silicide, aluminum oxide, silicon oxide, boronoxide, tantalum oxide, hafnium oxide, zirconium oxide, and titaniumoxide.
 10. The ceramic matrix composite tubular shell structure recitedin claim 1, wherein said ceramic matrix material comprises a singlephase of uniform material composition.
 11. The ceramic matrix compositetubular shell structure recited in claim 1, wherein said ceramic matrixmaterial comprises a multilayered phase including two or morealternating matrix layers having two or more ceramic matrix materialcompositions.
 12. The ceramic matrix composite tubular shell structurerecited in claim 1, wherein said ceramic matrix material comprises asingle phase of mixed material composition.
 13. The ceramic matrixcomposite tubular shell structure recited in claim 1, wherein saidtubular shell structure has a tubular geometry with said plurality ofcooling channels annularly arranged around and formed between said innerwall and said outer wall thereof.
 14. The ceramic matrix compositetubular shell structure recited in claim 1, wherein said tubular shellstructure is a cylindrical heat exchanger tube with said plurality ofcooling channels having a corresponding cylindrical profile and formedbetween said inner wall and said outer wall of said heat exchanger tube.15. The ceramic matrix composite tubular shell structure recited inclaim 1, wherein said tubular shell structure comprises a rocketpropulsion thrust chamber having a converging-diverging geometricprofile with said plurality of cooling channels having a correspondingconverging-diverging geometric profile and formed between said innerwall and said outer wall of said rocket propulsion thrust chamber. 16.The ceramic matrix composite tubular shell structure recited in claim 1,wherein said tubular shell structure has a longitudinal axis, and saidplurality of cooling channels formed between said inner wall and saidouter wall thereof are oriented axially with respect to the longitudinalaxis of said tubular shell structure.
 17. The ceramic matrix compositetubular shell structure recited in claim 1, wherein said tubular shellstructure has a longitudinal axis, and said plurality of coolingchannels formed between said inner wall and said outer wall thereof areoriented helically with respect to the longitudinal axis of said tubularshell structure.
 18. The ceramic matrix composite tubular shellstructure recited in claim 1, wherein said tubular shell structure has alongitudinal axis, and said plurality of cooling channels formed betweensaid inner wall and said outer wall thereof are oriented in parallelalignment and undulate sinusoidally with respect to the longitudinalaxis of said tubular shell structure.
 19. The ceramic matrix compositetubular shell structure recited in claim 1, wherein said plurality ofcooling channels are located in an annular array between said inner walland said outer wall of said tubular shell structure with each coolingchannel of said plurality thereof having a trapezoidal-shapedcross-sectional geometry.
 20. The ceramic matrix composite tubular shellstructure recited in claim 19, wherein said plurality of coolingchannels are nested in an annular assemblage between said inner wall andsaid outer wall to form said annular array thereof, said annular arrayof cooling channels located in intimate contact with said inner wall andsaid outer wall of said tubular shell structure.
 21. The ceramic matrixcomposite tubular shell structure recited in claim 1, wherein saidplurality of cooling channels form a corresponding plurality of radialwebs by which to mechanically couple said inner wall and said outer wallof said tubular shell into a high-efficiency unitized monocoquestructure.
 22. A method for manufacturing a ceramic matrix compositetubular shell structure having an inner wall, an outer wall and aplurality of cooling channel passages formed between said inner wall andsaid outer wall, wherein said ceramic matrix composite tubular shellstructure is manufactured by the steps comprising: a. fabricating aninner tooling mandrel for defining the inside geometry of said innerwall; b. fabricating a plurality of cooling channel tooling mandrelsegments for defining the inside geometry of said plurality of coolingchannel passages of said tubular shell structure; c. fabricating afibrous preform of refractory reinforcing fibers; d. depositing a fibercoating onto said fibrous preform which fully encapsulates therefractory reinforcing fibers thereof; e. depositing a ceramic matrixmaterial onto said fibrous preform to fully encapsulate and consolidatesaid coated fibrous preform into a densified fiber reinforced ceramicmatrix composite material; f. removing said inner tooling mandrel andsaid plurality of cooling channel tooling mandrel segments from saiddensified ceramic matrix composite tubular shell structure; g. machiningsaid densified ceramic matrix composite tubular shell structure; and h.depositing a ceramic seal coating material onto said densified ceramicmatrix composite tubular shell structure following said machining ofsaid tubular shell structure.
 23. The method for manufacturing a ceramicmatrix composite tubular shell structure recited in claim 22, whereinthe said inner tooling mandrel and said plurality of cooling channeltooling mandrel segments are fabricated from a material selected from agroup of materials comprising graphite, molybdenum and steel.
 24. Themethod for manufacturing a ceramic matrix composite tubular shellstructure recited in claim 22, wherein said inner tooling mandrel has ageometric profile.
 25. The method for manufacturing a ceramic matrixcomposite tubular shell structure recited in claim 24, wherein saidceramic matrix composite tubular shell structure comprises anactively-cooled heat exchanger tube having a cylindrical shape and saidinner tooling mandrel has a corresponding cylindrical geometric profile.26. The method for manufacturing a ceramic matrix composite tubularshell structure recited in claim 24, wherein said ceramic matrixcomposite tubular shell structure comprises an actively-cooled rocketpropulsion thrust chamber having a converging-diverging shape and saidinner tooling mandrel has a corresponding converging-diverging geometricprofile.
 27. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 26, wherein said inner toolingmandrel having said converging-diverging geometric profile comprises a2-piece assembly which is split at the location of minimum diameter andsecured by a bolted fastener.
 28. The method for manufacturing a ceramicmatrix composite tubular shell structure recited in claim 22, whereinsaid plurality of cooling channel tooling mandrel segments arefabricated by the steps comprising: a. forming a base tubular toolingmandrel having generally independent inner and outer geometric profilescorresponding to the cross-sectional dimensions of the plurality ofcooling passages; and b. sectioning said base tubular tooling mandrel atperiodic angular intervals to produce said plurality of cooling channeltooling mandrel segments.
 29. The method for manufacturing a ceramicmatrix composite tubular shell structure recited in claim 28, whereinthe inner geometric profile of said base tubular tooling mandrelconforms to the geometric profile of said inner tooling mandrel andincorporates a radial gap to accommodate the thickness of said fibrouspreform.
 30. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 28, wherein said plurality ofcooling channel tooling mandrel segments are fabricated by theadditional step of cutting said base tubular tooling mandrel at periodicintervals with a kerf width corresponding to the thickness of saidfibrous preform of refractory reinforcing fibers.
 31. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 28, wherein the step of sectioning said base tubular toolingmandrel includes the additional step of cutting said base tubulartooling mandrel at periodic intervals to produce a plurality of likeaxially extending cooling channel tooling mandrel segments.
 32. Themethod for manufacturing a ceramic matrix composite tubular shellstructure recited in claim 28, wherein the step of sectioning said basetubular tooling mandrel includes the additional step of cutting saidbase tubular tooling mandrel at periodic intervals to produce aplurality of like helically spiraling cooling channel tooling mandrelsegments.
 33. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 28, wherein the step ofsectioning said base tubular tooling mandrel includes the additionalstep of cutting said base tubular tooling mandrel at periodic intervalsto produce a plurality of like sinusoidally undulating cooling channeltooling mandrel segments.
 34. The method for manufacturing a ceramicmatrix composite tubular shell structure recited in claim 22, whereinsaid fibrous preform is fabricated from an assemblage of individualfiber preform elements produced by a textile fabrication processselected from a group of fabrication processes comprising weaving,braiding, knitting, fiber placement, filament winding, felting, andneedling.
 35. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 22, comprising the additionalstep of forming an inner wall fiber preform element onto said innertooling mandrel, wherein said inner wall fiber preform element isproduced by a textile fabrication process selected from a group offabrication processes comprising weaving, braiding, knitting, fiberplacement, filament winding, felting, and needling.
 36. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 22, comprising the additional step of forming a plurality ofcooling channel fiber preform elements onto repective ones of saidplurality of cooling channel tooling mandrel segments, wherein saidplurality of cooling channel fiber preform elements are produced by atextile fabrication process selected from a group of fabricationprocesses comprising weaving, braiding, knitting, fiber placement,filament winding, felting, and needling.
 37. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 36, wherein said plurality of cooling channel fiber preformelements are disposed around said inner wall fiber preform element in anannular array.
 38. The method for manufacturing a ceramic matrixcomposite tubular shell structure recited in claim 37, wherein saidplurality of cooling channel fiber preform elements disposed around saidinner wall fiber preform element in an annular array are compactedfirmly against said inner wall fiber preform element and secured by atight wrapping of tape to thereby form a fibrous preform sub-assembly.39. The method for manufacturing a ceramic matrix composite tubularshell structure recited in claim 22, comprising the additional step offorming an outer wall fiber preform element onto said fibrous preformsub-assembly to fabricate said fibrous preform, wherein said outer wallfiber preform element is produced by a textile fabrication processselected from a group of fabrication processes comprising weaving,braiding, knitting, fiber placement, filament winding, felting, andneedling.
 40. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 22, wherein the fiber coatingdeposited on said fibrous preform has a thickness of 0.05-5.0micrometers and is produced by a process selected from a group of fibercoating processes comprising chemical vapor infiltration (CVI), polymerprecursor impregnation/pyrolysis (PIP), reaction formation, andcombinations thereof.
 41. The method for manufacturing a ceramic matrixcomposite tubular shell structure recited in claim 40, wherein the fibercoating deposited on said fibrous preform is carbon produced by chemicalvapor infiltration using a carbon-forming precursor selected from agroup of chemical precursors comprising methane, propane, propylene, andmixtures thereof, which is pyrolytically decomposed into carbon at anelevated temperature of 950-1250° C. and at a reduced pressure of 1-250Torr.
 42. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 40, wherein the fiber coatingdeposited on said fibrous preform is boron nitride produced by chemicalvapor infiltration using a boron nitride-forming precursor selected froma group of chemical precursors comprising boron trichloride, borontriflouride, diborane, and mixtures thereof, which is reduced with areductant selected from a group of chemical reductants comprisingnitrogen, hydrogen, ammonia, and mixtures thereof to form boron nitrideat an elevated temperature of 700-1200° C. and at a reduced pressure of1-250 Torr.
 43. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 40, wherein said fibrouspreform is first coated with carbon followed by a boron carbide coatingproduced by chemical vapor infiltration using a boron carbide-formingprecursor selected from a group of chemical precursors comprising borontrichloride, boron triflouride, diborane, and mixtures thereof, which isreacted with a carbon-forming precursor selected from a group ofchemical reactants comprising methane, propane, propylene, and mixturesthereof to form boron carbide at an elevated temperature of 800-1100° C.and at a reduced pressure of 1-250 Torr.
 44. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 22, wherein said fiber coated fibrous preform is consolidatedwith a ceramic matrix material produced by a process selected from agroup of matrix consolidation processes comprising chemical vaporinfiltration (CVI), polymer precursor impregnation/pyrolysis (PIP), meltinfiltration (MI), reaction formation, and combinations thereof, whichfully encapsulates said coated reinforcing fibers of said fibrouspreform for transforming said fibrous preform into a dense, ceramicmatrix composite structure.
 45. The method for manufacturing a ceramicmatrix composite tubular shell structure recited in claim 44, whereinsaid ceramic matrix material is silicon carbide produced by chemicalvapor infiltration using a silicon carbide-forming precursor selectedfrom a group of chemical precursors comprising methyltrichlorosilane,dimethyldichlorosilane, silicon tetrachloride with methane, and mixturesthereof, which is reacted to form silicon carbide at an elevatedtemperature of 850-1150° C. and at a reduced pressure of 1-250 Torr. 46.The method for manufacturing a ceramic matrix composite tubular shellstructure recited in claim 44, wherein said ceramic matrix material iscarbon produced by chemical vapor infiltration using a carbon-formingprecursor selected from a group of chemical precursors comprisingmethane, propane, propylene, and mixtures thereof, which ispyrolytically decomposed into carbon at an elevated temperature of950-1250° C. and at a reduced pressure of 1-250 Torr.
 47. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 44, wherein said ceramic matrix material is boron carbideproduced by chemical vapor infiltration using a boron carbide-formingprecursor selected from a group of chemical precursors comprising borontrichloride, boron triflouride, diborane, and mixtures thereof, which isreacted with a carbon-forming precursor selected from a group ofchemical reactants comprising methane, propane, propylene, and mixturesthereof to form boron carbide at an elevated temperature of 800-1100° C.and at a reduced pressure of 1-250 Torr.
 48. The method formanufacturing a ceramic matrix composite tubular shell structure recitedin claim 22, wherein said plurality of cooling channel tooling mandrelsegments which define the inside geometry of said plurality of coolingchannel passages are destructively removed from said ceramic matrixcomposite tubular shell structure by a process selected from a group ofremoval processes comprising acid digestion, oxidation and gritblasting.
 49. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 48, wherein said plurality ofcooling channel tooling mandrel segments are fabricated from graphiteand are destructively removed from said ceramic matrix composite tubularshell structure by a process selected from a group of removal processescomprising oxidation and grit blasting.
 50. The method for manufacturinga ceramic matrix composite tubular shell structure recited in claim 48,wherein said plurality of cooling channel tooling mandrel segments arefabricated from metal and are destructively removed from said ceramicmatrix composite tubular shell structure by a process including aciddigestion.
 51. The method for manufacturing a ceramic matrix compositetubular shell structure recited in claim 22, wherein said densifiedfiber reinforced ceramic matrix composite tubular shell structure isseal coated with said ceramic seal coating material following removal ofsaid inner tooling mandrel and said plurality of cooling channel toolingmandrel segments from said ceramic matrix composite tubular shellstructure, said ceramic seal coating material produced by a processselected from a group of coating processes comprising chemical vaporinfiltration (CVI), polymer precursor impregnation/pyrolysis (PIP), meltinfiltration (MI), reaction formation, and combinations thereof.
 52. Themethod for manufacturing a ceramic matrix composite tubular shellstructure recited in claim 51, wherein said densified fiber reinforcedceramic matrix composite tubular shell structure is seal coated withsilicon carbide produced by chemical vapor infiltration using a siliconcarbide-forming precursor selected from a group of chemical precursorscomprising methyltrichlorosilane, dimethyldichlorosilane, silicontetrachloride and methane, which is reacted to form silicon carbide atan elevated temperature of 850-1150° C. and at a reduced pressure of1-250 Torr.